Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 670 AIRFOIL (goe670-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 670 AIRFOIL (goe670-il)
Reynolds number: 100,000
Max Cl/Cd: 54.43 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe670-il-100000-n5.txt
Download as CSV file: xf-goe670-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 670 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3084   0.10518   0.10028  -0.0360   1.0000   0.0728
  -9.750  -0.3230   0.10270   0.09786  -0.0372   1.0000   0.0737
  -9.250  -0.3931   0.10063   0.09544  -0.0364   1.0000   0.0598
  -9.000  -0.3921   0.09759   0.09243  -0.0355   1.0000   0.0578
  -8.750  -0.3978   0.09449   0.08938  -0.0350   1.0000   0.0559
  -8.250  -0.4647   0.08579   0.08093  -0.0369   1.0000   0.0497
  -8.000  -0.4722   0.08290   0.07808  -0.0358   1.0000   0.0495
  -7.750  -0.4809   0.07964   0.07485  -0.0352   1.0000   0.0494
  -7.500  -0.4910   0.07554   0.07074  -0.0357   1.0000   0.0496
  -7.000  -0.4867   0.07165   0.06688  -0.0329   0.9993   0.0520
  -6.750  -0.4641   0.06750   0.06263  -0.0377   0.9942   0.0533
  -6.500  -0.4452   0.06059   0.05549  -0.0445   0.9878   0.0529
  -6.250  -0.4263   0.05421   0.04880  -0.0495   0.9822   0.0527
  -6.000  -0.4054   0.04902   0.04330  -0.0528   0.9772   0.0531
  -5.750  -0.3833   0.04566   0.03968  -0.0547   0.9724   0.0549
  -5.500  -0.3604   0.04150   0.03513  -0.0567   0.9675   0.0566
  -5.250  -0.3375   0.03721   0.03035  -0.0581   0.9628   0.0569
  -5.000  -0.3146   0.03365   0.02626  -0.0587   0.9577   0.0576
  -4.750  -0.2862   0.03057   0.02261  -0.0598   0.9543   0.0587
  -4.500  -0.2633   0.02836   0.01988  -0.0594   0.9488   0.0599
  -4.250  -0.2342   0.02668   0.01778  -0.0600   0.9447   0.0623
  -4.000  -0.2017   0.02581   0.01682  -0.0613   0.9414   0.0645
  -3.750  -0.1774   0.02487   0.01570  -0.0608   0.9349   0.0662
  -3.500  -0.1450   0.02378   0.01438  -0.0618   0.9311   0.0684
  -3.250  -0.1139   0.02280   0.01318  -0.0624   0.9270   0.0710
  -3.000  -0.0872   0.02198   0.01214  -0.0621   0.9210   0.0738
  -2.750  -0.0538   0.02133   0.01150  -0.0633   0.9171   0.0773
  -2.500  -0.0248   0.02096   0.01110  -0.0635   0.9114   0.0832
  -2.250   0.0045   0.02045   0.01053  -0.0638   0.9054   0.0898
  -2.000   0.0394   0.01993   0.01001  -0.0651   0.9018   0.0974
  -1.750   0.0649   0.01953   0.00960  -0.0646   0.8948   0.1057
  -1.500   0.0964   0.01920   0.00923  -0.0653   0.8893   0.1180
  -1.250   0.1325   0.01871   0.00881  -0.0668   0.8859   0.1314
  -1.000   0.1554   0.01844   0.00861  -0.0659   0.8772   0.1444
  -0.750   0.1892   0.01806   0.00829  -0.0669   0.8725   0.1600
  -0.500   0.2213   0.01772   0.00805  -0.0677   0.8672   0.1798
  -0.250   0.2486   0.01743   0.00791  -0.0675   0.8592   0.2069
   0.000   0.2857   0.01687   0.00766  -0.0693   0.8552   0.2845
   0.500   0.4199   0.01467   0.00722  -0.0841   0.8427   1.0000
   0.750   0.4487   0.01443   0.00686  -0.0837   0.8280   1.0000
   1.000   0.4764   0.01423   0.00656  -0.0830   0.8126   1.0000
   1.250   0.5035   0.01407   0.00631  -0.0823   0.7963   1.0000
   1.500   0.5298   0.01397   0.00614  -0.0815   0.7803   1.0000
   1.750   0.5554   0.01394   0.00604  -0.0807   0.7652   1.0000
   2.000   0.5810   0.01393   0.00598  -0.0799   0.7504   1.0000
   2.250   0.6070   0.01391   0.00592  -0.0792   0.7344   1.0000
   2.500   0.6338   0.01390   0.00585  -0.0786   0.7174   1.0000
   2.750   0.6593   0.01394   0.00583  -0.0778   0.6994   1.0000
   3.000   0.6859   0.01398   0.00582  -0.0771   0.6805   1.0000
   3.250   0.7128   0.01406   0.00581  -0.0765   0.6602   1.0000
   3.500   0.7385   0.01419   0.00585  -0.0757   0.6389   1.0000
   3.750   0.7622   0.01439   0.00598  -0.0746   0.6163   1.0000
   4.000   0.7851   0.01462   0.00614  -0.0734   0.5928   1.0000
   4.250   0.8055   0.01488   0.00637  -0.0717   0.5650   1.0000
   4.500   0.8246   0.01515   0.00659  -0.0698   0.5296   1.0000
   4.750   0.8426   0.01550   0.00679  -0.0676   0.4829   1.0000
   5.000   0.8589   0.01598   0.00700  -0.0652   0.4362   1.0000
   5.250   0.8752   0.01656   0.00734  -0.0630   0.3985   1.0000
   5.500   0.8923   0.01714   0.00777  -0.0610   0.3686   1.0000
   5.750   0.9099   0.01773   0.00824  -0.0591   0.3441   1.0000
   6.000   0.9283   0.01828   0.00872  -0.0574   0.3249   1.0000
   6.250   0.9471   0.01883   0.00922  -0.0558   0.3090   1.0000
   6.500   0.9665   0.01936   0.00974  -0.0543   0.2962   1.0000
   6.750   0.9859   0.01992   0.01029  -0.0529   0.2852   1.0000
   7.000   1.0059   0.02045   0.01086  -0.0516   0.2750   1.0000
   7.250   1.0260   0.02099   0.01145  -0.0502   0.2651   1.0000
   7.500   1.0442   0.02157   0.01203  -0.0487   0.2530   1.0000
   7.750   1.0604   0.02212   0.01260  -0.0468   0.2374   1.0000
   8.000   1.0756   0.02265   0.01318  -0.0448   0.2201   1.0000
   8.250   1.0908   0.02318   0.01377  -0.0428   0.2030   1.0000
   8.500   1.1060   0.02370   0.01436  -0.0408   0.1862   1.0000
   8.750   1.1191   0.02425   0.01494  -0.0385   0.1637   1.0000
   9.000   1.1316   0.02493   0.01558  -0.0362   0.1274   1.0000
   9.250   1.1360   0.02636   0.01659  -0.0331   0.0847   1.0000
   9.500   1.1423   0.02790   0.01802  -0.0303   0.0577   1.0000
   9.750   1.1491   0.02938   0.01942  -0.0276   0.0464   1.0000
  10.000   1.1542   0.03091   0.02092  -0.0248   0.0413   1.0000
  10.250   1.1604   0.03230   0.02246  -0.0222   0.0377   1.0000
  10.500   1.1638   0.03389   0.02415  -0.0195   0.0349   1.0000
  10.750   1.1630   0.03578   0.02613  -0.0167   0.0332   1.0000
  11.000   1.1651   0.03751   0.02802  -0.0143   0.0320   1.0000
  11.250   1.1666   0.03937   0.03005  -0.0120   0.0308   1.0000
  11.500   1.1676   0.04134   0.03218  -0.0100   0.0296   1.0000
  11.750   1.1677   0.04348   0.03448  -0.0082   0.0286   1.0000
  12.000   1.1670   0.04577   0.03688  -0.0066   0.0274   1.0000
  12.250   1.1654   0.04828   0.03947  -0.0053   0.0265   1.0000
  12.500   1.1630   0.05106   0.04231  -0.0040   0.0256   1.0000
  12.750   1.1656   0.05349   0.04488  -0.0030   0.0251   1.0000
  13.000   1.1683   0.05605   0.04762  -0.0021   0.0244   1.0000
  13.250   1.1709   0.05870   0.05046  -0.0014   0.0240   1.0000
  13.500   1.1726   0.06156   0.05349  -0.0009   0.0235   1.0000
  13.750   1.1724   0.06468   0.05681  -0.0006   0.0231   1.0000
  14.000   1.1693   0.06818   0.06051  -0.0006   0.0224   1.0000
  14.250   1.1656   0.07184   0.06437  -0.0010   0.0220   1.0000
  14.500   1.1594   0.07595   0.06867  -0.0017   0.0216   1.0000
  14.750   1.1519   0.08036   0.07331  -0.0029   0.0213   1.0000
  15.000   1.1430   0.08511   0.07825  -0.0045   0.0210   1.0000
  15.250   1.1327   0.09035   0.08368  -0.0066   0.0208   1.0000
  15.500   1.1214   0.09599   0.08949  -0.0092   0.0205   1.0000
  15.750   1.1086   0.10218   0.09587  -0.0123   0.0203   1.0000
  16.000   1.0937   0.10915   0.10303  -0.0161   0.0203   1.0000
  16.250   1.0763   0.11715   0.11126  -0.0209   0.0204   1.0000
  16.500   1.0552   0.12651   0.12079  -0.0267   0.0205   1.0000
  16.750   1.0245   0.13925   0.13378  -0.0348   0.0212   1.0000
  17.000   0.9705   0.16084   0.15558  -0.0479   0.0229   1.0000
<< Back to GOE 670 AIRFOIL (goe670-il)

Polar data table (+)

Polar graphs


<< Back to GOE 670 AIRFOIL (goe670-il)